This invention relates to control surfaces for aerospacecraft, and more particularly to a ruddervator for an aerospacecraft incorporating a single piece, temperature resistant ceramic matrix composite shell secured over a composite structural member, wherein the structural member is adapted to be secured to a control element of the aerospacecraft.
Current control surfaces for advanced aerospacecraft are formed by a carbon-based ceramic matrix composite (CMC) hot structure with conventional rib-stiffened structure and a mechanically fastened skin. The X-37 aerospacecraft presently in use uses a control surface termed a xe2x80x9cruddervatorxe2x80x9d with the above-described construction, and makes use of carbon/silicon carbide (C/SiC). This construction is shown in FIG. 1. The mechanically fastened upper skin 10 is secured by a high temperature metal, ceramic or ceramic composite fasteners at locations 12 to an integral C/SiC lower skin and substructure 14. A C/SiC tail tip 16 is used to close the end of the ruddervator. A titanium spindle 16 is used to rotate the ruddervator as needed. Thermal protection system seals 18, 20 and ring 22 are used to help mount the ruddervator to the fuselage of the aerospacecraft.
The X-37 ruddervator approach described above uses an expensive 2800xc2x0 F. CMC system in a 2400xc2x0 F. xe2x80x9chot structurexe2x80x9d application and uses an aircraft-like structural approach at the elevated temperature. The term xe2x80x9chot structurexe2x80x9d refers to the temperature of the primary load-carrying structure, in this case the CMC and supports used at 2400xc2x0 F. This construction reduces the service life of the fasteners. Additionally, carbon-based CMCs generally require complex and costly tooling, unique and expensive infiltration/furnace facilities, and fabrication cycles of six months or more. The use of new materials under development, such as oxide fibers/oxide-matrix based CMC (oxide-CMC), provide opportunities to design control surfaces in novel and more cost-effective ways including, but not limited to, maintaining internal supports and attachments below 600xc2x0 F.
For present and planned reusable hypersonic vehicles there are also size constraints on control surfaces due to available volume which restrict the use of conventional, lower cost structure insulated with bonded tile thermal protection. The current solution is to use the CMC for control surface hot structure in areas which do not require their extreme high temperature properties. The result is high initial and recurring costs for these parts as well as weight penalties and high part counts. Without an order of magnitude reduction in thermal structure costs, commercial reusable access to space will be difficult, if not impossible, to achieve.
It is therefore a principal object of the present invention to provide a new construction for a ruddervator for an aerospacecraft which can be produced more inexpensively from a simpler fabrication process, and which has improved life and reliability over the conventional mechanically fastened upper skin-to-substructure approach presently in use for ruddervator applications.
It is another object of the present invention to provide a hybrid control surface for an aerospacecraft which can be manufactured more economically, which is simpler to repair, and which does not make use of typical mechanical fasteners to secure an upper skin to a substructure.
It is still another object of the present invention to provide a ruddervator for an aerospacecraft having a simplified design which requires significantly fewer independent component parts being needed in the construction of the ruddervator.
It is a further object of the present invention to provide a ruddervator for an aerospacecraft which can be constructed even more cost effectively, and which is reusable.
It is still another object of the present invention to provide a ruddervator for an aerospacecraft wherein the ruddervator employs a one piece, highly temperature resistant outer shell which is bonded to a composite structural member to provide a highly temperature resistant, lightweight and yet easy to manufacture assembly.
The above and other objects are provided by an airfoil in accordance with preferred embodiments of the present invention. The airfoil is specifically adapted to withstand the high temperatures encountered during hypersonic flight and is particularly suited for use as a ruddervator on an aerospacecraft.
The airfoil is comprised of a temperature resistant, ceramic matrix composite shell having an opening at one end and a hollowed out interior area. A structural member is inserted into the hollowed out interior area and bonded to an interior surface of the shell to form a structurally rigid airfoil assembly. A transition structure is secured to the structural member for interfacing the airfoil assembly to a control element of the space vehicle to permit the airfoil assembly to be controlled by the control element.
In one preferred embodiment the shell is comprised of a one-piece, oxide/oxide-based ceramic matrix composite (Oxide-CMC) shell. The structural member comprises a graphite composite structure having a graphite composite face sheet and a honeycomb core element.
During manufacture, the structural member is inserted into the hollowed out opening of the Oxide-CMC shell and is bonded thereto. The transition structure may be secured to the structural member either after the structural member is inserted into the shell or before attachment of the structural member to the shell. Finally, the transition structure is secured to a fuselage actuator spindle of the aerospacecraft.
The airfoil apparatus of the present invention thus forms a highly temperature resistant, easy to manufacture assembly. The assembly further reduces the cost and weight over present day ruddervator designs as a result of reducing the total number of parts required to form the ruddervator, in addition to providing a higher specific strength and stiffness of the materials used with the present invention.